FIG. 1 of the accompanying drawings is a schematic representation of a known aircraft ducted fan gas turbine engine 10 comprising, in axial flow series: an air intake 12, a propulsive fan 14 having a plurality of fan blades 16, an intermediate pressure compressor 18, a high-pressure compressor 20, a combustor 22, a high-pressure turbine 24, an intermediate pressure turbine 26, a low-pressure turbine 28 and a core exhaust nozzle 30. A nacelle 32 generally surrounds the engine 10 and defines the intake 12, a bypass duct 34 and a bypass exhaust nozzle 36. Electrical power for the aero engine and aircraft systems is generated by a wound field synchronous generator 38. The generator 38 is driven via a mechanical drive train 40 which includes an angle drive shaft 42, a step-aside gearbox 44 and a radial drive 46 which is coupled to the high pressure compressor 34 via a geared arrangement.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct 34 and exits the bypass exhaust nozzle 36 to provide the majority of the propulsive thrust produced by the engine 10. However, a proportion of the bypass flow is taken off and fed internally to various downstream (hot) portions of the engine to provide a flow of relatively cool air at locations or to components as or where necessary. The core flow enters, in axial flow series, the intermediate pressure compressor 18, high pressure compressor 20 and the combustor 22, where fuel is added to the compressed air and the mixture burnt. The hot combustion gas products expand through and drive the sequential high 24, intermediate 26, and low-pressure 28 turbines before being exhausted through the nozzle 30 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 24, 26, 28 respectively drive the high and intermediate pressure compressors 20, 18 and the fan 14 by interconnecting shafts 38, 40, 42.
In each of the turbine sections 24, 26, 28 the distance between the tips of the turbine blades and the radially inner surface of the turbine casing (or, more usually, the radially inner surface of the turbine blade track liner segments carried radially inwardly of, or forming part of, the casing) is known as the tip clearance. It is desirable for the tips of the turbine blades to rotate as close as possible to the engine casing without rubbing (or re-rubbing, in instances where it may be desirable to permit an initial or temporary degree of rubbing), because as the tip clearance increases, a portion of the expanded gas flow will pass through the tip clearance gap, and as a result the efficiency of the turbine decreases. This is known as over-tip leakage. The efficiency of the turbine, which partially depends upon tip clearance, directly affects the specific fuel consumption (SFC) of the engine. Accordingly, as tip clearance increases, SFC also rises, which is disadvantageous.
Under conditions of transient increases in engine power, such as during take-off or a step climb of an aircraft, as the disc and the blades of the turbine rotate, centrifugal force and increasing thermal loads cause the disc and blades to expand in a radial direction. The turbine casing also expands as it heats up, but typically there is a mismatch in radial expansion between the disc/blades and the casing. Specifically, the blades will normally expand radially more quickly than the casing, thereby reducing the blade tip clearance and potentially leading to rubbing (or re-rubbing) as the tips of blades come into contact with the interior of the casing, until the casing itself heats up and expands sufficiently to increase the tip clearance again back to an optimum distance. To accommodate such behavior, working tip clearances may thus need to be over-compensated for, leading to tip clearances under stable engine power conditions being greater than optimum for a major part of any flight profile or cycle.
In an effort to alleviate such a disadvantage, there have been several proposals in recent years which involve actively controlling the temperature of the turbine casing to a desired degree as or when required, so that the radial expansion of the casing can be more accurately matched in a responsive manner to that of the turbine disc/blades at any point or stage in a flight cycle, even under conditions of especially enhanced engine power such as a step climb.
One such known system is that disclosed in EP2372105A, which is shown schematically for a typical HP turbine architecture, by way of example, in FIG. 2 of the accompanying drawings. Here the proposed system allows a typical additional blade tip running gap associated with step climbs, being an excess over the optimum tip clearance gap G, to be removed, by ensuring that the casing can be thermally expanded very quickly in the event of a step climb. For this purpose a discrete, thin impingement plate 50 formed with any suitable pattern of impingement through-holes 52 therein is located radially within the casing 60 above (i.e. radially outwardly of) the carrier 70 and blade track liner segment 80 carried thereon. The principle is that in the event of a step climb, a valve 90 above (i.e. radially outwardly of) the casing 60 is opened, and air is drawn through the impingement holes 52 in the impingement plate 50 to heat and therefore thermally expand the casing 60 in a short space of time. It should be noted that before the valve 90 opens, the casing 60 is typically cooler than one or more of the higher stages of compressor air on account of the external cooling from the outboard bypass air. In this manner a more responsive arrangement for heating (and cooling, if required) the turbine casing to control the tip clearance of the rotating turbine blades at any given stage of a flight profile, e.g. even upon a step climb, is provided. This makes it possible to maintain a minimal tip clearance whilst preventing rubbing (or re-rubbing) of the blades against the turbine casing during transient increases in engine power, while maintaining a relatively high level of engine efficiency during stable cruise conditions.
In practical forms this known system shown in FIG. 2 typically employs a thin tinware sheet as the discrete impingement plate 50, which is not only very difficult to assemble, but also leads to significant problems in terms of air sealing and position control, since thin continuous sheet material typically has a much quicker thermal reaction time than the engine casing material itself, which may lead to buckling and thus making the impingement distance between it and the casing much harder to control for optimum impingement performance. Leakage around the impingement plate 50, leading to compromised engine efficiency, may also be a practical problem.
Another, similar, known system for actively controlling the temperature of the turbine casing is that disclosed in EP2546471A. In this system a dedicated inboard duct is provided, adjacent an inboard surface of the turbine casing, which has an outboard facing wall with a plurality of impingement holes formed therein and opening towards the inboard surface of the casing, through which impingement holes temperature control fluid can pass from within the inboard duct to impinge upon the inboard surface of the casing to regulate its temperature. The temperature control fluid, e.g. air from a compressor stage of the engine or even air taken from two or more locations at different temperatures so as to be mixed to a desired optimum temperature, may be re-circulated internally.
A disadvantage of this known system, however, is that the dedicated inboard duct is constituted by an additional component that adds weight, cost and build complexity to the overall arrangement. It also means that the recirculating temperature control air is applied to the casing substantially continuously, thereby requiring substantially constant temperature control regardless of whether a specific casing temperature requirement, e.g. heating during a step climb, is actually required in any given stage of an overall flight profile.
It is therefore an object of the present invention to provide a constructionally simpler, cheaper and more efficient system for actively controlling the temperature of the turbine casing of a gas turbine engine, especially for improving the responsiveness of an arrangement for heating and/or cooling a turbine casing to more efficiently control turbine blade tip clearance during transient increases in engine power during a flight profile, e.g. during step climbs.